Gas turbine power plant having means to cool and means to compress combustion products passing through the turbine



Dec. 30, 1952 R BlRMANN 2,623,357 GAS TURBINE POWER PLANT HAVING MEANS TO COOL AND MEANS TO COMPRESS COMBUSTION PROiUCTS PASSING THROUGH THE TURBINE Filed Sept. 6, 1945 11 Sheets-Sheet l mnt/Ess" INVENTOR.

DGC- 30, l952 R. BIRMANN 2,623,357

GAS TURBINE LPGNER PLANT HAVING MEANS Tov Cool.. yAND :MEANS To COMPRESS COMBUSTION PRODUCT-s PASSING THROUGH THE TURBINE Filed Sept. 6, 1945 vll Sheets-Sheet 2 sus]` y y v fragt/5X6.

R. BIRMANN 11 Sheets-Sheet 3 GAS TURBINE POWER PLANT HAVING MEANS TO COOL. AND MEANS TO COMPRESS COMBUSTION PRODUCTS PASSING THROUGH THE TURBINE Filed Sept. 6, 1945 Dec. 30, 1952 INVENTOR Fao/UZ z 'rma/.272

ffrJv/-yfx Dec. 30, 1952 R. BIRMANN 2,523,357

GAS TURBINE POWER PLANT HAVING MEANS TO COOL AND MEANS TO COMFRESS COMBUSTION PRODUCTS PASSING THROUGH THE TURBINE Filed Sept. 6, 1945 1l Sheets-Sheet 4 A /Zf F76. /ZZ

Dec. 30, 1952 R BIRMANN GAS TURBINE POWER PLANT lHAVING MEANS TO COOL AND MEANS TO COMPRESS COMBUSTION PROLUCTS PASSING THROUGH THE TURBINE l1 Sheets-Sheet 5 Filed Sept. 6, 1945 Dec. 30, 1952 R BIRMANN 2,623,357

GAS TURBINE POWER PLANT HAVING MEANS TO COOL AND MEANS TO COMPRESS COMBUSTION PROE'UCTS PASSING THROUGHTHE TURBINE Filed Sept. 6, 1945 11 Sheets-Sheet 6 INVENIGR.

Dec. 30, 1952 R. EMRMANN 2,623,357

GAS TUEBINE POWER PLANT HAVING MEANS TO GOOI. ANO MEANS TO COMPRESS COMEUSTION PROSUCTS PASSING THROUGH THE TURBINE Filed Sept. 6, 1945 1l Sheets-Sheet '7 INVENTOE gli y gz/vays;

Dec. 30, 1952 R..B1RMANN 2,623,357

GAS TURBINE POWER PLANT HAVING MEANS TO COOL. AND MEANS TO COMPRESS COMBLJ'STTGN.` PROUCTS PASSING THROUGH THE TURBINE Filed Sept. 6, 1945 ll Sheets-Sheet 8 Dec. 30, 1952 R. BIRMANN GAS TURBINE POWER PLANT HAVING MEANS TO COOL AND MEANS TO COMPRESS COMBUSTION PRODUCTS PASSING THROUGH THE TURBINE Filed Sept. 6, 1945 WIWESS' l1 Sheets-Sheet 9` www-0R Dec. 30, 1952 R, BlRMAN 2,623,357

GAS TURBINE POWER PLANT HAVING MEA TO COOL AND MEANS TO COMPRESS COMBUSTION PRODUCTS PASSING THROUGH THE TURBINE Filed Sept. 6r 1945 11 Sheets-Sheet 10` R. BlRMANN 2,623,357

1l Sheets-Sheet ll Dec. 30, 1952 GAS TURBINE POWER PLANT HAVING MEANS TO COOL AND MEANS TO COMPRESS COMBUSTION FROLUCTS PASSING THROUGH THE TURBINE Flled Sept. 6, 1945 lllllflll IllIlI-lllllllllilh gia/@4.1

Patented Dec. 30, 1952 UNITED STATES PATENT OFFICE GAS TURBINE POWER PLANT HAVING MEANS TO COOL AND MEANS TO COM- PRESS COMBUSTION PRODUCTS PASS- ING THROUGH THE TURBIN E 13 Claims.

This invention relates to gas turbine power plants. Particularly the invention has reference to a power plant utilizing a novel thermodynamic cycle and to the elements of such power plant whereby high eiciency of the cycle is achieved.

One object of the invention relates to the provision of the new cycle. This, in contrast to cycles heretofore known, involves the utilization of considerably less air per net brake horsepower output, with the result that the heat energy introduced is more eciently converted into mechanical, and in certain cases additional jet energy, by minimizing the air flow proportionately used, the quantity of which air flow for a given power output involves proportional losses in a compressor and turbine even though a large compressor and a large turbine may be as ellicient or more efficient than small ones. Further, the thermodynamic eiciency is greatly improved in that expansion of the gases takes place through unusually great temperature and pressure drops. Other features of this cycle will become apparent hereafter. l

Another object of the invention relates to a novel turbine, not only, in its preferred form, capable of efficiently utilizing extremely high velocity flow of driving gases, but of operating with a pressure rise through, at least, several stages thereof whereby certain important features of the improved cycle are realized, though the turbine is of more general use. The turbine further embodies certain structural features, cooling provisions, etc. which constitute additional objects of the invention.

Another object of the invention relates to the provision of improved heat eXchangeRrneans for the cooling of high temperature gasesflowing at extremely high velocities, the cooling taking place with a minimum loss of kinetic energy.

These and other objects of the invention particularly relating to details will become apparent from the following description read in"conjunction with the accompanying drawings in which:

Figures 1A to 1D, inclusive, are successive views partly in axial section of the parts of the improved power plant taken in order from its forward to its rearward end;

Figure 2 is a side elevation of the compressor showing in particular its various connections to an intercooler and its air delivery connections to the air heating means;

Figure 3 is an end view of the compressor unit looking in the direction of the arrow 3 in Figure 2;

Figure 4 is a cylindrical development of a section taken on a surface of revolution immediately inside the casing, looking inwardly between the arrows 4 4 in Figure 1B, and showing in particun lar the stationary and rotating vanes of the rst compressor stage;

Figure 5 is a developed view of a section taken on the surface of revolution indicated at 5-5 in Figure 1C showing in particular the air passagesfrom the compressor to the heat exchanger and the arrangement of fins therein;

Figure 6 is` a section taken on the plane indicated at 6 6 of Figure 1C;`

Figure 7 is a developed view of a section taken on the surface of revolution indicated at 'I-T of Figures 1C and 1D and showing in part the arrangement of the combustion chambers, nozzles, A.

the whirl chamber, and the moving and stationary vanes of the turbine;

Figure 8 is an enlarged sectional view of the the plant;

'Figure ll is a fragmentary section taken on the radial plane indicated at I I-I I in Figure 10;

Figure l2 is a development on the section taken on the cylindrical surface indicated at IZ-IZ in Figures 10 and l1;

Figure 13 is a temperature-entropy diagram explanatory ofl the improved thermodynamic cycle; and

Figure 14 is a diagrammatic illustration of the application of the invention to a low power turbine plant operating at pressures below atmospheric pressure.-

There will first be described the general mechanical features of the improved power plant, following which the design of certain elements and the operation will be described.

Generally stated, the power plant comprises a turbine connected to drive both a compressor and a propeller, `the compressor furnishing the air for combustion of fuel to provide the turbine driving gases.

The shaft 4 of the propeller spinner 2 which carries the propeller blades is driven through reduction gearing 5 by the shaft 8 connected through flexible coupling I0 to the shaft I2 mounted in a bearing I3, which shaft I2 serves as the forward mounting of the compressor rotor. The rear mounting of the compressor rotor is through the medium of a shaft I4 supported by 3 a bearing I6. A flexible coupling I8 connects shaft I4 With the shaft 20 of the turbine, which shaft is mounted in a bearing 22. The turbine at its rear end is provided with a shaft extension 24 mounted in a bearing 26.

Air is led to the compressor through the passage 28, which, in the case of aircraft applications, may receive its. air through wing slots or other openings directed forwardly to take advantage of the ram eifect of movement through the air, Whi-ch thereby gives rise to Ian initial pressure and/ or velocity of approach to the compresns-or. Immediately in front of the first' compressor stage, station-ary vanes 30 are provid-ed to impart spin to the air for proper reception by the compressor. The compressor operates on the principle of vortex ow and in such fashion Ithat 'at its inlet at any radius the peripheral component of the absolute velocity lof the air is inversely propcrtional to the radius. The vanes 3U are accordingly designed in the same general fashion as the. spin-imparting vanes at the inlets of conventional axial How compressors to produce a vortex iiow having'the characteristics just mentioned. TheVv proportionality constant of this flow is the same or substantially the same for rated conditions as the' proportionality constant at the inlet of the rst stageimpeller. In this fashion, sm'ooth en trance, without loss, into the impeller'is assured. The vanes 3l) should be designed to be of airf'oil cross-section in accordance with good 'aerodynamic principles to secure spin of the air Without burbling and without impact losses.

The compressor illustrated comprises ve stages. The rst three Aare in direct series" and located back to back' with respect to the last Itwo which are als'o in direct'series with each other'. Except for' the' radial dimensions of the passages land the vane' heights, the various compressor' stages `are identical and in particular the impeller hubs vare identical, except for' the outlet portions of the third and fifth stages. The hubs comprise an nul-ar shells 32 and 34, providing, as the inner boundaries of the air passages, surf-aces of revolution. These shells are Welded together as indicated -at 33 and in series with each other as indicated at d2. In ord-er to impart the necessary mechanic-al strength, each hub section is provided with a dished disc element 36 or 38 Welded' to 'the peripheral portions of the hub shell. of the-se dis-c elements is such as to secure maximum tensile strength and rigidity against distortion at high speeds yof operation. in accordance w'ith'this, these :disc elements are approximately ofa shape which Would naturally be assumed by a iiexible loaded surface having junction to an axis of rotation at spaced points and subjected to centrifugal forces, i. e., a shape which might' be described as a three dimensional catenary, the axial section `of which approximates a parabola or catenary. In order to secure rigidity -against axial displacement cylindrical struts 31 join the' disc elements 36 and 38.

The joined hub sections of the third and fifth stages are provided with shaft-like extensions Welded toge'ther'as indicated at 44 and are surrounded by a seal 222 of a type hereafter described Which maintains separation of the outlet pressureregions oi the thirdi'and iifth stages.

As Will become apparent hereafter, the vanes of i each impeller could theoretically be continuous from their inlet to their outlet portions. However, to secure the proper fio-vv characteristics this would mean that except for some parti-cular radial section, they would. deviate very` substan The shape Y 4 tially .from a radial direction and hence be subject to ver;7 large deflecting stresses during high speed operation. Accordingly it is desirable to break up the vanes into a series of groups of axially shorter vanes as indicated in Figures 1B and 4. A series lof vanes 4B is thus provided at the inlet, a second series 48 is providedtokdene the intermediate portions cf the flow passages, anda third series 5G is provided at the outlet. When vanes of lthis character are used it becomes possible to Y provide in the c-ase of each vane a strictly radial section infsome particular radial plane, while the remaining portions of the vane fore and aft of this radial section dc not deviate too greatly from a radial condition and will thus be subject only to rlow non-destructive stresses. In general, it may be said that the vanes lie along theoretical skew surf-aces and are cf airf-oil sections by annular surfaces taken in the direction cf the flow stream.

The general appearance of these vanes will. be evident from the cylindrical development illustrated in Figure e. l They `are such asto maintain substantially lthedretically proper oW in any radial plane, being designed 'to' secure the proper absorption `of energy by the airfrom the rotating vanes.

The housing for the impeller comprises a portion 52 enclosing, with suitable clearance, the rotating vanes and a second portion 5d, which carries the spin-reducing vanes 5E', Which latter vanes in turn' have a suitable clearance with" the hub. The vanes 56 are similar to the ia'- peller vanes in the sense that they are designed to maintain substantially theoretically proper fi-ow at any particular section, butV are so dee signed that the vspin component along any streamline decreases in the direction through' their passages.4 Thev4 outlet ends of the impeller vanes and the inlet ends ofthe spin-reducing vanes are such as to give to the air flow the 'A same proportionality constant of vortex flow so that the iiow from-the rotating' irnpeller vanes to the stationary spin-reducing vanes is smooth. In the case of the spin-reducing vanes, it is not necessary to take into account centrifugal stresses and consequently each vane may be continuous from its inlet to its outlet. in any casa'l the spin-reducing vanes,'both when single and when multiple, should' desirably have air-foilV characteristics in the direction of now;

The description, given with respect to the first stage impeller, applies equally to the other rimpellers, the `vanevgroups of Whichjare indi-v cated at 58, 62, e2 yand $6; Likewise, the de?" sc'ription given for the-spin-'reducing vanes' 'is applicable' to' `the similar vanes of. the' second and'fcurth stages indicated at Gil' and. ed. Thesole difference' involved, as maybe -noted'fro rn the drawing', is that the'annular' areasof the:

succesive compressors, i. e., the vaneheights,

are reduced to take into consideration the compression which occurs and which results in greater density of the owing air and, consequently, for the saine velocity, a greater mass,

loW per unitgannular area;

In the case of the third stage impeller,

air is to pass tc an interccoler 'and consequently it is not necessary to cause the air te ilovv radial- 1y inwardly to the inlet of the next stage'. Ac-

fleet it in a radial direction into the annular passage 66.

l Referring to Figure 2, the air from the passage 65 is taken off tangentially through the connections 68` and 'Hl to the inlets 12 and 14 of intercoolers 16 and 18 of conventional aircooled type, in which intercoolers it flows through suitable tubing in heat exchange with air taken in through openings in the aircraft wings or other openings and discharged from the intercooler to the atmosphere. The compressed air leaves the intercoolers at 8|! and B2 and is delivered through the tangential connections 84 and 86 into the annular passage 88. From this annular passage the air flows through varies 90 arranged to impart to it a vortex type of flow for proper impactless approach to the rotating vanes Q2 of the fourth stage compressor. The fourth stage impeller delivers the air through the spin-reducing vanes 94 to the fth l I, the air flow is divided into a series of passages |Il2 each of which is in turn subdivided into an outer passage H36 and an inner passage |08 (see Figures 1C and 5). spectively communicate with outer and inner pre-heating chambers HD and ||2. rIhe outer pre-heating chamber is bounded on its inner side by the casing ||1 of an annular gas flow passage and by the casing IIE! of the turbine. the outerwall of the passage ||El being cylindrical. Extending axially through the chamber I5 and arranged in axial planes are a large number of ns H4, between whichthe air flows. These fins are metallically bonded to the casings ||1 and H8 so as to provide maximum heat conduction from these casings through the fins to the air. The compressed air flowing through the passage lll] thus leaves at highly heated.

The passage ||2 is somewhat similar in that it is furnished with fins ||5 similar to the fins I 4, and having metallic connections to the inner cylindrical wall H9 of the combustion gas passage. passage I 0, the air is substantially heated.

The air flowing from the passages I IG and I2 then flows forwardly through the passages |22 and |25, which are insulated as indicated at |24 to prevent loss of heat. These passages |22 and, |28 communicate with a common annular passage |26, which is closed except for a series of inlets individual to a series of combustion chambers and controlled by adjustable valves |32. The series of combustion chambers is provided about the circumference of the plant and comprises outer spaces |54 communicating with the inlets |30 into which spaces the air flows to pass inwardly to support combustion through the end opening |36 and a series of openings |38 provided by the nested corrugated cylindrical members |45, constituting a flame chamber, surrounding the axis along which fuel is projected from the fuel atomizing nozzle |42 (ignition being started in conventional fashion, not illustrated). This arrangement is such that the These passages re- 5 5 Accordingly in this passage as in the.

gradual in passing between the vanos. Vregions are so shaped not only to provide mechanical strength, in that by their use the tur- 6. spray of fuel is burned progressively along vits length by annular streams of air so that complete combustion occurs before there is reached 4the region |44, which is lined with ceramic Ina-- Aterial to withstand the intense heat. The mem` bers |45 are formed of heat-resisting alloy andv Aare protected from the intense heat of the flame by reason of the annular flow of air passing Athrough the spaces |38. The region |44 of each of the combustion chambers communicates` with an individual ceramic nozzle |46, of some high- 4'ly refractory material such as fused alumina or sidably at an angle to the circumferential direction of about 16, so as to circulate therethrough in'Y advancing spiral fashion in their approach to the rst series of turbine vanes in such fashion that the streamlines of flow maire about one to one and a quarter complete revoltuions about the axis between the nozzles and the first set ofturbine vanes. A greater number of revolutions may be made for more effective cooling, but

at the expense of greater friction loss.

The inner surfaces of this whirl chamber are highly polished for smoothness and are oxidized or `have their heat absorption increased in. some other manner for reasons hereafter stated.

`The turbine hub sections are indicated at |50, 52 and |54. These comprise shells formed as surfaces of revolution as indicated and carry the series of vanes |56, |58, and |62, the particular nature and operation of which will be de-y scribed hereafter. The spaces |64 and |66 be ktween the vanos |56 and |58 and between the vanes |58 and I6!) are annular in shape and as indicated have portions which, as viewed in Figures 1C and 1D, have less radii than the preceding and succeeding vanes. While this may seem to involve a rather sharp diversion of direction of now, it is to `be noted that the flow has a very high circumferential component and a relatively low axial component of flow so that, as a matterA of fact, the radial diversion of the now is very These bine hub portions |50, |52 and |54 may be shaped to secure the maximum mechanical 'strength necessary to withstand the centrifugal Yforces at the high temperatures which exist at the varies, but also for a thermodynamic advantage hereafter described. In the case of the flow between the vanes |60 and |52, however, the temperature has already been suflicien-tly reduced so that this same necessity for extreme strength no longer applies. Furthermore, the Whirl component is now quite low so that it would be difficult to cause the now to flow inwardly and then outwardly without separation from the annular walls. Consequently the space |68 may be substantially cylindrical.

From the vanes |62, the gases pass throughr the redirecting stationary blades |10 and finally pass through the buckets |12 of a pressure stage.

being discharged therefrom into the annular space |14 from which now occurs through the tail passage |15 to a nozzle |16, from which the gases may issueA to the atmosphere in the formA of-"arelatively high velocity jet giving rise to a jet propulsion thrust.-

, Within the hub portions of the turbine. there are; provided discs |18, |80., |812., |85, |86, |188 and |90: as. illustrated. These are so shaped (as described in connection with the impellers) andsecured together. as to provide the. maximum strength for the support of the shell sections of the. hub which are welded to these discs and to each. other.A Cylindrical struts |19 provide axial rigidity for the hub structure. The disc |92 carrying the buckets |12 is secured to the disc |90.

In order to provide cooling, the air, bled from passages |04 previously referred to, passesV at |94 aboutv the bearings |6 and 22 and the intermediate coupling |8, and thence into the passage |96 from which it ows through. slots |98 inthe member |78 into the space between discs l'land. |80. This air then passes through. the slots 202, 204, 208, 2|0 and 2|4 in succession, thereby-reaching the interiorA of the various cells inside the rotor. Finally,v it enters into the space 216 between discs |88 and |90 and passes therefrom at high temperature through nozzles 2|0, whichare. so directed as to act as backwardly discharging converging-diverging nozzles so that their discharge enters inner extensions 220 of the stationary blading |70 in smooth fashion at high velocity, the pressu-re drop through the nozzles being considerably in excess of the critical pressure. drop. This stationary blading acts to. provide redirecting passages to direct the air through theroot portions of the buckets |12 so that it aids the turbine rotation. It nally mingles withV the combustion gases in the annular region |14.

In order to secure successful cooling of the vanes which arewelded into slots in the hub shells-, there are. provided inside the hubs as integral extensions of the vanes extended surface ar-1 rangements in the form of ns 200, 22S, and 2|2 so'that heat may be effectively conducted from the vanes in direct fashion. These iins are milled in an angular direction so that smooth flow occurs through them, the air passing substantially axially relative to the rotorthrough the slots |98, 202, 204, 208 and 2|0 so that rela-- tive. rotary iiow occurs as the air ows outwardly from the slots into the. region of the ns.

At. various points in the power plant necessity arises for sealing the junction of relatively rotating parts against pressure differences. {Sealing' means isv provided at: such places asr indcated; at 222, 224, 226 and 220. There will now bel-described a typical.y sealing means which by modiflcationof form may be used at these various points. Specically, the sealing means at 2.24 will be described Vwith reference to Figures 10,v 11 and 12.

A stationary partition is indicated at 230. on the. left side of which a high pressure exists and on the right side of which a low pressure exists. The. shaft I4 is provided with a flange 232 having aradial plane face against which bears the ra.` dial plane face of a carbon ring 234, this ring being spaced from the shaft at 235, the dimension of which space is considerably exaggerated in Figure I0. Thev right hand face of the carbon ring is provided, with a series of groovesy as indicated ajt 238 extending in the direction ofilogarithmic spirals vand'communicating with wedge` like depressions or pockets 239, the Wedge angle of which' is very slight, for example, of the order of a quartervto a half degree. The ring 234 is mounted in a metal ring 240 which is provided with a conical'face242 opposed to a conical face 8,y 2.44.011- therpartitlonJ 230; An annular helical` spring (a so-called garter spring) 246 is tern sioned4 to; reactbetween these wedge-shaped surfaces so. as. to urge the member'240 and the. car-f bon ring 2'34to the right under slightv pressure. An annular groove 2.48 in the member 240 carries a series of piston rings 250 which bear against a cylindrical surface of the partition 230. uncle-r the pressure of a corrugated circular spring indicated` at. 252.

The arrangement of the elements just described is such that an air lm is maintained for lubrication purposes between the carbon ring and the flange 232. Air rotating at high velocity enters the grooves 233 aided by centrifugal forces and is then distributed into the wedge-shaped space 239, passing as a nlm from its trailing edge. By reason of the wedge effect this air Vilm is continuously maintained, being of a thickness of only a small fraction of a thousandth of an inch. The leakage from the inside to the outside of. the relatively-moving surfaces is accordingly veryy slight. In order to maintain thislm in a proper condition for lubrication it is necessary that the pressure of the carbon ring toward the ilangebe low and of predetermined value depending solely upon the spring force. This condition is achieved by the proper area relationships. The high pressure forces the assembly of the rings 234 and 220 to the right, At the same time a force isv exerted to the left by the pressure of the air nlm which has a gradient ranging from the high pressure at the inside, i. e.,the clearance 236, to the low pressure at the outside. By making the bearing area of the carbon ring of proper larger size"than thearea presented to the high pressure the arrangement is balanced to the extent of the pressure difference so that the bearing pressure is substantially solely due to the action of the spring 24B. In thisy fashion an effective seal is maintained with very little friction. The high speed of rotation provides an air film of excellent lubricating qualities despite the low viscosity of the air and the normal absence of any lubricating qualities of an air film. Only at very low speeds of operation, as, for example, in starting up the system, will there occur actual contact between the carbon ringV and the iiange.

As will be apparent hereafter, the invention contemplates for best operation an extremely high supersonic velocity of approach of the driv-l ing gases to the turbine. `It is obvious that conventional turbines are inadequate to meet this situation. Not only would theireiciency be-exe .tremely low in view of the high gas velocity as compared with the permissible peripheral velocity for entrance into conventional buckets, but an extreme pressure change would necessarily occur in inefficient fashion in conventional buckets, causing, in particular a temperature rise which would .exceed the permissible temperatureV for the buckets. In addition, there is desirably secured in the rotating turbine blading'a pressure 'rise from inlet to outlet as will be discussed hereafter in reference to the cycle. Conventional turbines operate with an overall pressure drop in the rotatingV blading having a limit of zero in the case of impulse turbines.

In the case vof the improved turbine the vanes |55 of the first setv are airfoils of the type designed for eicient operation at. supersonic velocities and are designed to reduce the velocity to sonic or slightly less than sonic velocity. In the case of a set of vanes such as at |50 designed Ito receive the air at supersonic velocities the passages between the vanesare very important. Deceleration is always connected with compression shock, but it has been found that with proper shape' and proportion of passages, deceleration from a Mach number of 1.5 to a Mach number of 1.0 may be very efficiently secured, e. g. with an efficiency of 93%. The vanes should be arranged as illustrated in Figure 7. rFhey are characterized by their provision of passages involving afrelatively slight contraction ratio of about 1.1 for a Mach number reduction from 1.5 .to 1, deflection' of the iiow from an average inlet angle of about 15, proper for reception of the spinning approaching' gas, to an average outlet angle of about 13, sharp leading and trailing edges, and relatively slight convexity with maximum thickness in the mid-chord vicinity, the location of the maximum thickness and the general blade section being determined from the point of view of theupassages and that of designing the vanes as compression shock control devices rather than from the point of view of supersonic airfoils.'

. In View of the substantial reduction of Mach number, there occurs therein a very considerable part of the total compression taking place inthe negative-reaction blading. Furthermore, these Janes contribute substantially to the mechanical workoutput of the turbine by reason of causing` a` deceleration of the supersonic flow rather than by reason of the deflection thereof.

Since the varies |56 receive gas having a vortex flow through the whirl chamber, they should have a twist so as to receive the vortex fiow smoothly and also to discharge the flow in the form vof a vortex for properly stable flow through the space H7 and to the second set of vanes.

The velocities of approach to the second set of vanes will be sonic or less, and accordingly these vanes |58 are designed for such velocities. They may take the form illustrated, having slender inlet sections with only slightly rounded inlet edges, and having either' a straight or slightly curved mean thickness line to secure a relatively slight deflection of theflow; e. g. for an average inlet angle of 13 to an average outlet angle of 17, and providing diverging passages, the equivalent cone divergence angle of which is of the order of 2 to 4. These vanes |58 are also twisted to receive and deliver vortex liow.

Usually the gases approaching the third set f varies will have substantially subsonic velocities and accordingly the airfoilvanes |60 may be designed to correspond thereto, with rounded entrance, quite substantial camber, and with maximum thickness slightly ahead of the center. The relative velocity of discharge may, in the case of these vanes, be almost axial. These vanes, also, are twisted to receive and deliver vortex flow. l

Y The fourth set of, vanes |62 are also of subsonic twisted type and serve to deflect the ow rearwardly relative to thejrotor to such extent that the total deflection from inlet to outlet of the turbine may be of the order of.100 or more. This high total deflection which could not possibly be secured efficiently by means of single buckets of a conventional turbine at the velocities involved can be achieved atvery high elficiency by the luse of the multiple airfoil cascade described.

It may be noted that while, as illustrated. the deflection is carried well beyond an axial direction of relative flow, this in` some cases. may

not be desirable inasmuch as less net pressure rise will occur. For extremely high initial temperature it may be desirable to secure the maximum degree of expansion through the nozzles, hence in such case the flow relative to the blading may not be carried beyond an axial direction.

Considering the turbine constituted by the blades |56 to |62, it will be evident that the substantially constant blade height which exists in the preferred form of the invention is necessarily tied up with such deceleration of the flow, without expansion, that the ow must take place with a pressure rise. This is in contrast to the reaction type of turbine which involves a positive pressure drop and to the impulse type of turbine involving a zero pressure drop. For convenience of reference, therefore, the present turbine will be designated by the apt term negative-reaction turbine with the understanding that this refers to a turbine such as that described involving a pressure rise, i. e., a compressor action.

It may be noted that constant blade height is preferred only from the point of View that the initial vortex flow is advantageously designed with the smallest possible helix angle, resulting in the maximum permissible vane height from the point of view of strength, i. e., the inlet angles of the varies should be minimum. If other considerations than strength were involved, the vane heights should increase towards the discharge.

It will be evident from the principles of thermodynamics that the compression occurring in the turbine must necessarily be accompanied by a rise of temperature. If, therefore, the gases leave the vanes |576 at the highest permissible temperature for the materials and construction thereof, it will be evident that the subsequent vanes would be heated above permissible temperatures unless provision was made for the removal of heat during the compression. As a matter of fact, the passage of the gases through f the turbine is accompanied by a very effective heat removal to the end that the process through the turbine may be kept essentially isothermal. The portions of the housing between the rows of vanes are desirably polished just as in the case of the Whirl passage |48 so that heat dissipation by vortex cooling takes place even though the circumferential velocity of flow decreases through the vanes. l

`In the vortex cooling in the whirl chamber and between the sets of vanes, the flow is subjected to a centrifugal eld which is of major value in securing good heat transfer. The centrifugal acceleration is of the order of 500,000 times the acceleration due to gravity. The extremely high velocities result in high Reynolds numbers and consequently high Nusselt numbers corresponding to high rates of heat conduction from the flow to the walls. The conditions for heat radiation from the gases to the walls are also ideal due to the existence of large gas masses ofconsiderable radial and axial depth, relatively low rate of axial movement of these gas masses, and excellent geometrical conditions for high radiation due to the curvature of the walls. and, in the case of the interblade spaces, due to the deep dips of the passages.

AThe centrifugal field `also contributes to a great degreeto the heat transfer.` Due to the high acceleration, convection forces of great l l magnitude are involved (as compared with the usual slight ones due to gravity alone). As an ultimate result heat transfer rates are substan tially increased Without resulting ina proportional increase of friction losses.

The conditions are also such to provide Aan ideal situation for radiation of heat from the' vanes. It will be noted in particular that in the case of the vanes |56 and |58, through which thev major slowing down of the now occurs and in which, accordingly, the maximum liberation of heat occurs, their aspect is such that, major proportions of the normals to the surface of any vane do not substantially meet any other vane so that radiant energy may pass therefrom directly to the quite cool housing of the turbine and to the cooled rotor surfaces. The dips in thehousing and rotor surfaces between the sets of vanes serve to receive this radiation and absorb it very effectively. For this purpose these surfaces are oxidized or'otherwise blackened or their heat absorption increased inV some other manner. Additionally, heat is also transferred by conduction from the vanes to the interior extended surface at 206), 206 and 2|'2. This eX- tended surface is very effectively cooled by the air within the turbine shell, the action Ybeing particularly augmented by the centrifugal effect. on the air which causes its normal flow to. take place toward the periphery with a relative circumferential velocity'. there existing here. also vortex cooling as described above. As a result of these very effective fashions in which heat is removed the compression which occurs is maintained from an overall standpoint substantiallyA isothermal.

The residual leaving velocity of the gases. from the vanos |62 is turned into use by being directed through the blad-ing |163 into the blading |12 of an additional turbine, the velocities now being sufiiciently low that a turbine of more nearly conventional type is satisfactory. The stationary blading Ill) is designed to receive and deliver vortex iioW, being twisted in a fashion which will be clear from the discussion of the compressor.- Since the discharge angle of the vanes |62 is large, little or no turning is requiredin the guide vanes |10,A this condition COntributing substantially to the efficiency. The blading |12 is twisted in. accordance with vortex requirements so that pure impulse action exists at the base of the blades with reaction of the order, of 50% at the blade tips. The final leaving velocity of the gases, now traveling with little, if any, peripheral component of flow, may be utilized directly inI the form of a jet, or if operationis carried out* under certainv conditions as, for example, at high altitude, a pressure drop through the outlet nozzle |16 may serve to give rise to further jet thrust.

The interior cooling air enters athigh pressure` the varies 2|8 which form converging-diverging nozzles toproduce` high velocity jets to create a s'ubistantial` reaction aiding rotation of the turbine. The redirecting blades 220 for this air direct it to Vthe, blades |72 of the final turbine, between the innermost sections of which blades this air flows to, aid in driving the same. VThe redirectiing nozzles provide vortexV flow so. that this prevails throughout the height of blades |12.

Having now described the turbine, the overall cycle of operation may be described.

The compression of air in the vfirst Vthree stages of the` compressor is substantially adiabatic, the deviation from adiabatic compression being due fir l2 to: losses in the compressor. Thisair is then cooled in the intercooler at approximatelyl con,-` stant pressure. The temperature of the compressed air isreduced in the intercooler to a temperature somewhat above the temperature of the atn'uispheric air which absorbs the heat of the compressed air. In the last two compressor stages there is again approximately adiabatic compression with the resulting attainment; of a final. quitey high pressure ratio `achieved in eiilcient fashion.`

The lheating-of the air which has been compressed now occurs in -two general stages. lThe rst stageis'involved in the passages I l0 and` I2 at approximately constant pressure. While the air has, compared with the combustion gases.

relatively low flow velocities, neverthelessythe heat transfer is very effectively carried out in View of the'extensive surface afforded by the fins ||4 and ||6 and the high density of the com.- pressed air.. This surface may be, for example. of theV order of fifteen to twenty times vthe area. of the surface exposed to the gases in the Whirl passage |48. The highly heated air then ows through the insulated passages |22 and |28 to the combustion chamber.

In the combustion chamber the air is further heated at substantially constant pressure to an extremely high temperature. The relation of the air to the fuel is such as to secure complete combustion with only such excess asl will keep the temperature at about 3500 absolute. While this is lower than the temperature which would result from an ideal fuel-air ratio', it is, nevertheless,

l extremely high and should be as high yas can'be permitted, consistent with the later attainment at the turbine `variesof atemperature which can l be tolerated bythe material thereof.

By the construction of the combustion chamber heretofore described it will be evident that though` the temperature of the gases leaving the. name may be extremely high, nevertheless, effec.- tive cooling of the name passage Walls ismain'f tained` to prevent their destruction by yreasor'iof the progressive addition of 'air in the formof annular increments. According-to the invention, therefore, an attempt is desirablymade to secure at the inlets to the nozzles |46 the maximum temperature permissible: in View of subsequent phases of the cycle. While excess air is added the quantity ladded is very much lessl than has been necessary in. previously known cycles to; avoid: destructive,- temperatur'es at., the turbine.

Ihe expansion in the nozzles |46 takes place substantially adiabatioally from the pressure` de.-V livered. by thecompressor to the pressureexisti-ng in the whirl passages |48'.- In the casev of a unit of the type herein illustrated, the pressure inthe Whirl chamber may bel about 0.85 of the atmos-l pheric pressure under sea level operating oonditions. Accordingly, what might be called superexpansion takes place with both the attainment of extremely high kinetic energy of the gas and with intense accompanying cooling. Howevenit- Willbeevident that., particularly where there are nov/eight andspace limitations of the, type; ex:-` isting.. in aircraft applications, higher compres` sion ratiosmay be desirablev to-securestill; higher thermal efficiency and in such cases; thevexpan sion. through the nozzles would only be carried. far enough to achieve suii'ciently'low tempera--v tures at the turbine vanes. The expansiom in other Words, neednot,l inV such. cases, be carriedz Y to subatmospheric pressures. Y TheV temperature of the gases issuing from the 13 nozzles will normally still be far above a safe operating temperature for vanes of the best known materials. Accordingly, there occurs intense further cooling in the whirl chamber |48. This cooling, however, taking place at substantially constant pressure, is not accompanied by any substantial reduction of kinetic energy. It is presently well established that the transfer of heat between a iiowing fluid and a solid wall increases With the relative velocity and that this still holds for extremely high velocities. In particular vortex cooling occurs with the advantages outlined above. An intense interchange of heat occurs from the whirling gases through the polished walls and fins to the air on the exterior of the Whirl passage. Not only does this occur by reason of the long helical Contact of the flowing gases with the walls of the whirl chamber but additionally by radiation to the walls, the conditions being ideal for such radiation. The resulting high transfer rate is sufficient to accomplish a lowering of temperature to a point at which the gases may safely be admitted to the turbine vanes. Even though thepassage |48 may be of relatively short axiaL length, nevertheless, the path of flow of the gases passing therethrough is quite long because of the low pitch of the whirl.

The cooling, it is to be noted, does not represent a loss of useful energy of the gases. The kinetic energy, except for low frictional losses, is maintained. The heat transferred to the air is not lost because this raises the air temperature prior to the burning of the fuel to -a quite high temperature so that the desired temperature at the inlets to the nozzles may be secured with correspending economy of fuel, i. e., given a particular desired temperature at the nozzle entrances the fuel required to reach this temperature is less to the extent of the preheating of the air which supports the combustion.

There next occurs the flow of gases through the turbine which, through the first three sets of vanes, is, from an overall standpoint, approximately isothermal. Actually, however, the phases involved are relatively complicated and are as follows:

In the first set of vanes |55 approximately adiabatic compression takes place with resulting rise in temperature to a maximum value approxi--` mating the limiting temperature at which the vanes may permissibly operate. This temperature is the temperature for which the preceding matters of design are aimed, i. e., the combustion, the expansion and the cooling are so carried out that this maximum temperature is not exceeded. The adiabatic compression takes place with accompanying reduction of the peripheral components of the absolute velocity of the gases so that in leaving the vanes the whirl is substantially less than at the entrance. The result is the imparting of torque to the vanes to contribute to the drive of the turbine. As pointed out previously, the gases both entering and leaving the vanes desirably have vortex flow.

VThe gases now now, still with a quite high whirl component, rst inwardly and then outwardly through the passage |64, being cooled as they do so. During the period of reduction of the mean radius of flow some expansion occurs and then with increase of radius of flow there occurs compression. The net result, however, as viewed from the outlet of the vanes |56 to the inlet of the vanes |58, is an approximate cooling at constant pressure, the pressures at these two end points being substantially the same. Due to the 14 high whirl component and radiation, there continues, however, the same general type of cooling as occurred 'in the whirl passage |48 so that at the entrance to the second set of vanes 58 the temperature is such that the compression occurring in these vanes will not result in the attainment of dangerous temperatures.

In passage through the vanes |58 essentially the same operations occur as through the vanes |56. Approximately adiabatic compression again oc curs, the whirl velocities are still further reduced to impart driving torque lto the vanes, and vortex flow is provided at the discharge. In the region |66 beyond the vanes |58 there is again expansion followed by compression accompanied by cooling in such fashion that at the entrance to the vanes |50 the pressure is approximately the same as at 'the exit from the vanes |58.

In the vanes |60 there is also approximately adiabatic compression, again with the attainment of a temperature 'which does not exceed a safe value. The region |68 into which the vanes |60 discharge with maintenance of vortex flow is of substantially constant mean radius so that the cooling which occurs in this passage is substantially at constant pressure without the expansion and compression previously referred to. The reason for this is that the temperature has now been so reduced that there is little point in securing the increase in velocity which has been achievedy by the expansion in the regions |64 and |55, which increase of velocity contributes substantially to the heat transfer.

In the fourth row of vanes |62 expansion occurs rather than compression. At the exit from these vanes the gases still possess considerable kinetic energy and redirection of the gases takes place in the xed blades |10, the gases being delivered to the vanes |12 of the nal stage. The gases leaving the final stage may pass directly to the atmosphere, or through the nozzle |16 in which a final pressure drop occurs giving rise to a propulsion jet.

The advantages of the foregoing cycle will be made clear by reference to the temperature-entropy diagram of Figure 13 which is typical for the type of cycle involved. Compression occurs in the low pressure compressor A to B, the atmospheric line being represented by AX. Cooling takes place from B to C. This is followed by compression in the high pressure stages of the compressor from C to D. The compressions are approximately adiabatic but not strictly so in view of losses in the compressor stages.

Heating of the air then occurs, first by heat exchange and then by combustion of the fuel to raise the temperature along the constant pressure curve DE to the point E at the temperature T. The initial part of this curve may be taken to represent the lower heat reservoir for a succeeding part p of the cycle which involves the cooling of the whirling gases issuing from the nozzles.

Following this the gases are cooled at substantially constant pressure in the whirl passage while their kinetic energy is maintained, cooling taking place along the subatmospheric pressure line FY to the point G.

In the rst row of vanes of the negative-reaction type compression takes place from the point G to the point H.

HJ represents the expansion and recompres-` 'atteste sion in the 'vortex space lbetween the 'nr-'st and second' rows of vanes. While this is taking place heat is continuously withdrawn and imparted partly to the compressed combustion air rlowing through the fins on the outside o'f the vortex space and partly to the precompressed cooling air flowing along the interior `sur-face of the vturbine -rotor. The compression which follows the initial expansion .in the vortex space accordingly takes Aplace at a lower ten'lperatr'e than 'the 'mean temperature Vprevailing du'l ng the'expansion. If no friction losses occurred this phase of the cycle, involving compression under lower meantemperature conditions than those prevailing during the expansion, would 'fresult in a net pressure increase. However, unavoidable friction losses enter into the pictilre so that actually the pressure increase to the point J may merely 'involve a nal pressureap- Dri''r'n'ately' that Of the point H asV indicated on the diagram.

In the second row of varies compression 'takes place from J to K ina fashion similar to that occurring between G and H. Then in the 'foie lowing vortex chamber there is again expansion followed by compression along the line winding up approximately at the -salnelpre'ssur at .L as at K.

In the third Vrow ofyanes compression takes place trom L to M. This is followed vby cooling from :M to N at substantially constant pressure.

--In the fourth lrow of vanes expansion takes place from N to O'. At this point G the gases still possess considerable kinetic energy which, in the entropy diagram, is represented by the adi'- abatic line OP. Finally PQ representsV the lipansi'on line to atmospheric pressure representing the change of state occurring in the iinal pressure stage. `In the event that a l'jet is provided this line also includes the change of 'state occurring in the jet. It may be noted that from G to P the voverall process is approximately isothermal, and for a simplified view of the cyclev it may be said that the cycle involves approximately isothermal compression through 'the first three sets of vanes. Y

The temperature T of the point Halt will be noted, is the highest Vtempe'rature involved in the cycle in the turbine biasing. Actually-,the turbine blading will not be subjected to an average temperature quite this high but rather"V 'a t m perature somewhere between the temperatures ofthe points G and H.

In comparison with this there may be Vcoinsideredthe Braytoncycle which corresponds to al blade temperature of T'. This Brayton eyl would be represented by AJSCDUVAQthe 'zigzag line UV representing the process through Vthe conventional 'type of turbine which, for the pressurerati'o involved, vwould necessarily be ofV multis-t'a'ge type. It will be 'notedthat the highest temperature involved at U must be much less than T. This requires in the conventional ilse of the Brayton cycle the introduction of sunicient excess air so that 'the temperature at theenf trance to the nozzles is not higher than thai-at U. In comparison with this it'will be seen Vthat despite the heating 'of 'the gases' to the mu'c'h higher temperature corresponding to E, never theless by virtue of the superexparision andthe cooling along the line FG, the temperature of the vanes of the first set is maintainedwithin the proper limit. At the saine time there is transformed into useful energy the heat Arepre'- sented by vthe quite large 'area by 'which the Y -1'6 'present cycl'e `area exceeds area. A

The temperature-entropy diagram, however, does not rtell the whole story inasmuch "as 'for a givenfa'ddiftion' 1of?. heat by the fuel, representing 'the solo energy input to the power plant, the pres t cycle requires the handling er 'a fmueh lower amount of air. In the case of the' Brayton 'Cy-Cle :between three and 'five times thaln'O-llt of air would have *te te handled. l

Tfli' cooling all flowing through ythe rotor -S-hll of the -tl'l'b involves -a Separate Icycle of its own as shown `lat 'the lower left in Figure 13. Compression Aof this cooling air occurs in the low `pres i Vstag-Gs of the compressor Vfrtlm vA. to B;

)the Brayton l cycle t *"s 'then cooled inthe i-n-tercooler from B to C, artes 'which fit ls heated the intel-ier ef the rotorreln C to -Rl It is "nally expanded along the line its the reac'tr nozzles located at v'2 l a 'and the Stationary voztssags v2.".1) and the blabllng' if thenal presslr'e ystage. vFor Cooling pl'lrposcs the air used ny be of 'the order 0":672'0721Y of 'the main new so that to determine lthe'ne't power output ci the cooling en in relation to the main ycle about In'ieelfth of the total area snonl-dn'bie compared with the total area. of the large diagram representing the main cycle.

The lc'li'rect power output obtainable ffrom the cooling air is, therefore, rather small. #However, indirectly, the cooling air maires another contribution to the power output of the unit which is more substantial, increasing the `area of the main cycle by `contributing to making nearly isothermal the process from G to P. Ii this latter process was not isothermal a corre*- spo'nding wedge would be deducted vfrom vthe main cycle las will 'be evident from the diagram. In addition, a lower ini-tial temperature would have 'to vbe used in order to arrive at the same blade temperatures, which would lower the'cycle ellci'ency.

From the foregoing it will be noted that the heat imparted to both the combustion air fand the internal 'cool'n'g'a'i of Vthe 'turbine is b'el'g L'Sed in a double fashol'whoh is Very important. The compressed combustion air and the cooling air eonstittea lower ht reservoir 'for theA cycle. Heat is ac'ccfrdingw 'withdrawn from the .cycle in such 'a manner that the conversion ler neat energy into mechanical energy isV increased. The lie'at witlfldrawnv from the cyole is then returned to the cycle Vin vsuch manner that less heat be added by fuel combustion, in the oase of 'the combustion air,r anda part of the 'heat withdrawn is converted into mechanical energy in the case ofthe 'coo-lingair. Y

While the present invention'has been described in its vpreferred (for-m, it will y'be evident that substantial advantages can be gained even -by modications thereof which do not attain the full valuable aspects or what is herein described.

For example, by the 'useof a negative-reaction turbine involving compression it would be possible to eliminate the cooling in the whi-r-l cha-mbe'r Md, directing the nozzle gases directly to thev turbine. This possibility exists if ahigher compression ratio -is involved through the negativereaction turbine tobring the output `pressure of the turbine to the initial pressure (e. g., atmospheric pressure) without the use of a turbinev stage involving a pressure drop. Under such circumstances the' supereexp'ansi'on alone may be ilse'd to'provide a su'ficient drop to a tempera# ture which can be tolerated by the tur-bine' varies.`

`17 This will be evident from the temperatureentropy diagram constituting Figure 13.

Heat recovery from the exhaust gases may be effected in substantially conventional fashion for the purpose of driving an independent turbine or alternatively a turbine of conventional type which may be geared or directly connected to the main turbine.

It will be evident that the invention is applicable to a closed cycle type of operation in which a mass of air or other gases is continuously recirculated at increased pressures, heat being added to this air thro-ugh heat exchange means without admixture with it of combustion gases. The heat in this case may be added by the combustion of fuel, electrically, or by the utilization of atomic energy.

Multiple staging of the system described, is of course, possible and advantageous in certain cases. For example, there may be cited the design of a gas turbine for a lower power output than that of the power plant described above, such as would be required for driving automotive vehicles, which design also illustrates the application of the invention to a system in which combustion or other heating of air takes place at atmospheric pressure and the cycle operates, in general, at subatmospheric pressures. For such an application a high starting torque is necessary (as compared with the plant heretofore described which may be started by auxiliary means, not shown, to give it an initial spin) which can only be obtained by separating the turbine' driving an exhauster-compressor from the turbine delivering the useful mechanical output. Such an arrangement, for example, would consist of a negative-reaction turbine, the nozzle of which would receive gases substantially at atmospheric pressure and would expand the gases to a subatmospheric pressure, e. g., 9.2 atmosphere, in a whirl chamber cooled by the incoming air in its approach to the combustion chamber, this air also cooling the negative-reaction turbine in the fashion previously described. This turbine would drive a compressor and would discharge into the power turbine, which, in automotive applications, would be geared to the wheels. This power turbine would expand from approximately atmospheric pressure at the inlet to the power turbine nozzles, achieved by the pressure rise in the negative-reaction turbine, to subatmospheric pressure (e. g., 0.3 atmosphere) at the discharge. The gases leaving the power turbine would be cooled in a conventional heat exchanger and would then be recompressed to atmospheric pressure by the aforementioned exhauster-compressor. The power turbine may thus` operate over a wide speed range starting from a standstill while the exhauster-compressor could operate continuously at full speed. The exhauster-compressor may be of the -type of compressor herein described. The power turbine may be a conventional impulse or reaction turbine.

A power plant of the foregoing type is diagrammed in Figure 14. Air enters at 260 and passes through the heat exchange passages 262 and 264 of the type previously described and thence by Way of the return passage 266 into the combustion chambers' 268 into which fuel is injected at 2li). The combustion products then expand through nozzles 212 into the whirl chamber 213 from which they flow through the vanes 2M, 276 and V21S of the negative-reaction turbine. The gases exhausted fromthis turbine,now,at approximately atmospheric pressure due to the compressor action, flow through the passage 235) and are expanded in the nozzles 282 cooperating with the blading 284 of the power turbine 25e. This turbine, through reduction gearing 233, drives the power output shaft 290. The gases exhausted from the turbine 286 are then cooled in the exhaust gas cooler 29,2 in which the cooling medium may be water lor air. The cool gases then flow through the passage 294 to the multiple stage compressor 286 of the type heretofore described, being discharged through the diffuser 258 to the atmosphere. As previously described, subatmospheric pressures exist in the Whirl chamber and also on the exhaust side of the turbine 284. It will be evident that the mechanicalv construction, not shown in detail, is essentially similar to that of the aircraft powerplant heretofore described in detail, with such modifications, as, for example, decrease of the number of nozzles, as are consistent with the lower` power output. Various matters described herein are claimed in my applications Serial Numbers 38,995 and 73,977, filed, respectively, July 16, 1948 and February l, 1949.

In View of the foregoing and other obvious modifications of the invention, it will be understood that the following claims should not be construed as limited to the particular preferred design herein described which is `particularly adapted for aircraft use;

What I claim and desire to protect by Letters Patent is:

1. In combination, means for supplying high temperature elastic fluid, a turbine having rotating blading, a stationary nozzle receiving elastic fluid from said supplying means, expanding said fluid, and discharging said fluid at high velocity, means comprising solely stationary guiding walls for leading said fluid at substantially said high velocity from said nozzle to said turbine blading through an extended path longerthan the diameter of the turbine blading, and means for leading elastic fluid in heat exchange relationship with the rst mentioned fluid leading means and thence to said supplying means to effect substantial heat exchange between i the elastic fluid streams flowing through the respective leading means.

2. In combination, means for supplying high temperature elastic fluid, a turbine having rotating blading, a stationary nozzle receiving elastic fluid from said supplying means, expanding said fluid, and discharging said fluid at high velocity, means comprising solely stationary guiding Walls for leading said fluid at substantially saidhigh velocity from said nozzle to said turbine blading through an extended path longer than the diameter of the turbine blading, and means for leading elastic fluid in heat exchange relationship with the first mentioned fluid leading means and thence to said supplying means to effect substantial heat exchange between the elastic fluid streams flowing through the respective leading means, said turbine having at least one compressor stage so that flow occurs therein with a rise of pressure of the rst mentioned fluid.

3. In combination, means for supplying high temperature elastic fluid, a turbine, an annular whirl chambena nozzlereceiving elastic fluid from said supplying means and discharging said fluid in skew direction into said whirl chamber, the nozzle being of a type effecting expansion and high velocity of discharge of the elastic fluidso thatl the discharged fluid has a helical path in the whirl chamberandflows therein at high velocity, said whirl chamber being constructed and 

